1. Field of the Invention
The present invention relates generally to an industrial gas turbine engine, and more specifically to a cooling and sealing arrangement for a first stage blade outer air seal and second vane airfoil.
2. Description of the Related Art including information disclosed under 37 CFR 1.97 and 1.98
In a gas turbine engine, a compressor delivers compressed air into a combustor to be burned with a fuel and produce a hot gas flow, which is then passed through a turbine to drive the compressor and, in an industrial gas turbine (IGT) engine, also to drive an electric generator.
An industrial gas turbine engine has several major design differences from an aero engine which is used to power an aircraft. An IGT is much larger and is designed to operate for long periods of time between shut-downs, typically in the range of 48,000 hours. In the turbine section, stages of rotor blades rotate within the engine. The rotor blade tips form a gap with an outer shroud assembly to limit the amount of hot gas flow leakage through this gap. The outer shrouds are formed of segments that are secured in place to the engine casing. FIG. 1 shows a prior art arrangement. The shroud segment 11 is secured between two isolation rings 13 and 14 that extend from the engine casing 15. A gap in which the hot gas can enter is formed between the isolation ring 14 and the adjacent vane shroud 16 in which the vane airfoil 17 extends. An aero engine does not use these isolation rings because of the heavy weight. In the aero engine, the outer shrouds are secured directly to the engine casing. As such, the stator vanes in the aero engine cannot be removed from the engine without disassembling the entire turbine section.
In the IGT, the isolation rings allow for the stator vanes to be removed without having to disassemble the rest of the turbine. The stator vanes can be removed radially outward without removing the outer shroud segments. Because of this arrangement in the IGT outer shroud and isolation ring structure, a gap is formed between the shroud segments and the adjacent stator vane assembly. FIG. 1 shows this prior art turbine inter-stage structure for the first stage blade outer air seal (BOAS) and second vane airfoil. In this type of blade outer air seal with downstream vane component design, there is no sealing or cooling arrangement to prevent the hot gas ingression along the axial gap. As a result, hot gas flows in and out along the inter-stage gaps and an over-temperature occurs at the blade outer air seal edges and the blade isolation ring corresponding to the hot gas injection location. this over-temperature issue is more pronounced when a lack of adequate inter-stage gap purge air for the axial gap exists in combination with a strong bow wave induced by the low solidity vane airfoil creates a high circumferential pressure variation that will force the hot gas into the inter-stage gap.
A prior art reference U.S. Pat. No. 4,177,044 issued to Riedmiller et al on Dec. 4, 1979 and entitled COMBINED TURBINE SHROUD AND VANE SUPPORT STRUCTURE discloses an aero gas turbine engine with a shroud segment forming a blade tip gap with the first stage turbine blades in which the shroud segment includes a plurality of axial holes (66 in the Riedmiller patent) that open into a shroud rear groove and a plurality of axial slots formed underneath an inner flange extending into the shroud rear groove from the vane assembly. Cooling air flows into an impingement cavity of the outer shroud and then through the axial holes and axial slots and into the gap to prevent hot gas from injecting into the gap. The main difference between the Riedmiller patent and the present invention is that the Riedmiller structure is for an aero engine and not an IGT. In the aero engine of the Riedmiller patent, the adjacent vane cannot be removed in the radial direction because of the inner flange extending into the shroud rear groove. Thus, the adjacent vane assembly cannot be removed from the engine without disassembling the shroud segments as well. Also, the Riedmiller patent lacks the use of the isolation rings to support the shroud segments.